This invention relates generally to gas turbine components, and more particularly to turbine airfoils.
A gas turbine engine includes a compressor that provides pressurized air to a combustor wherein the air is mixed with fuel and ignited for generating hot combustion gases. These gases flow downstream to one or more turbines that extract energy therefrom to power the compressor and provide useful work such as powering an aircraft in flight. In a turbofan engine, which typically includes a fan placed at the front of the core engine, a high pressure turbine powers the compressor of the core engine. A low pressure turbine is disposed downstream from the high pressure turbine for powering the fan. Each turbine stage commonly includes a stationary turbine nozzle followed by a turbine rotor.
Each turbine rotor carries a circumferential array of airfoil-shaped turbine blades adapted to extract energy from the combustion gases exiting the core. These blades are typically constructed by casting from high-temperature resistant alloys (e.g. “superalloys”. The first rotor stage, immediately downstream of the combustor, is usually internally cooled and has a hollow interior with one or more serpentine passages, film cooling holes, trailing edge slots or holes, and the like. The subsequent rotor stages are not subject to the extreme high temperature of the first stage and thus does not require cooling. To reduce the weight of the later-stage airfoils, they often include a hollowed-out internal portion referred to as a “weight reduction plenum”. Unfortunately, the shape of the prior art weight reduction plenum can cause the airfoil to fail and separate at the interface between the hollow plenum and the solid airfoil portion.
Accordingly, there is a need for a turbine blade having low weight while maintaining high strength.